Vented tangential on-board injector for a gas turbine engine

ABSTRACT

An on-board injector that delivers discharge air toward a turbine rotor of a gas turbine engine includes a second wall spaced form a first wall to define an annular inlet about an engine longitudinal axis and a multiple of airfoil shapes between the first wall and the second wall to segregate discharge air from the annular inlet, and a multiple of bypass apertures each along a radial axis transverse to the engine longitudinal axis through each of the multiple of airfoil shapes and the respective first wall, the second wall.

CROSS REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.15/913,269, filed Mar. 6, 2018, which is a divisional of U.S. patentapplication Ser. No. 14/609,926, filed Jan. 30, 2015, now U.S. Pat. No.9,945,248, issued Apr. 17, 2018, which claims the benefit of provisionalapplication Ser. No. 61/973,338, filed Apr. 1, 2014, which are alsoincorporated herein by reference.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This disclosure was made with Government support awarded by The UnitedStates Air Force. The Government has certain rights in this disclosure.

BACKGROUND

The present disclosure relates to a gas turbine engine and, moreparticularly, to Tangential On-Board Injectors.

Gas turbine engines, such as those that power modern commercial andmilitary aircraft, generally include a compressor section to pressurizean airflow, a combustor section to burn a hydrocarbon fuel in thepresence of the pressurized air, and a turbine section to extract energyfrom the resultant combustion gases. The hot gases expanded within theturbine section produce a gas stream across alternating rows ofstationary turbine stator vanes and rotating turbine rotor bladesproduce power.

Internal secondary flow systems transfer cooling air that bypasses thecombustor section to a turbine rotor assembly for subsequentdistribution to the interior of the rotor blades through a tangentialon-board injector (TOBI). Accelerating the cooling air through a nozzle,and swirling the air with the rotation of the turbine rotor, reduces thetemperature of the cooling air as it is injected on board the turbinerotor.

The volume and direction of the cooling air are features of thesecondary flow system effectiveness and overall engine performance. Thesecondary flow system should provide a desired metered amount of coolingair as additional cooling air will penalize efficiency of the engine,while too little cooling air may result in overheating of the rotatingturbine disks, blades, and seals Additionally, the secondary flow systemdirects purge air within the engine to prevent hot gas ingestion in theturbine rim cavities. Typically, rotating Knife Edge (K/E) seals, inconjunction with honeycomb seal lands, are used to control the amount ofpurge mass flow needed to seal and purge cavities. Other seals such asbrush seals and contact seals can be used for this purpose with varyingsealing effectiveness; however a certain amount of purge mass flow isrequired to properly protect the turbine rotor from hot-gas ingestion atthe rim cavities. Heat pickup due to passage heat conduction/convection,rotor cooling, and windage losses due to the rotation effects of thedisks and rotating seals, increases the temperature of the purge flow asit passes through the engine. It is desirable to use this heated purgeair to satisfy the rim cavity mass flow requirement, as its coolingeffectiveness has been greatly reduced and no longer has ability to dofurther rotor/blade cooling.

The temperature of blade cooling air is negatively affected by theundesirable mixing of the cooling air with the purge air, which is airthat flows past the various seals and cavities within the gas turbineengine towards the TOBI. When air exits the TOBI, the flow does notpurely flow into the rotor/blade as rotor cavity purge air must flowacross the TOBI discharge stream. The crossing flows mix, and pollutesthe TOBI flow. The net result is the air flowing to the blade may berelatively hotter and thereby relatively less thermally efficient.

SUMMARY

An on-board injector that delivers discharge air toward a turbine rotorof a gas turbine engine according to one disclosed non-limitingembodiment of the present disclosure includes a first wall; a secondwall spaced from the first wall to define an annular inlet about anengine axis; and a multiple of airfoil shapes between the first wall andthe second wall to segregate discharge air from the annular inlet, and amultiple of bypass apertures each along an axis transverse to the engineaxis through each of the multiple of airfoil shapes and the respectivefirst and second wall.

A further embodiment of the present disclosure includes, wherein themultiple of airfoil shapes include a trailing edge arranged about 80degrees to an engine axis.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the multiple of airfoil shapes include atrailing edge arranged about 10 degrees to circumferential.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the multiple of airfoil shapes define acascade exit to segregate the discharge air.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein each the multiple of airfoil shapes includea pressure side and a suction side, the pressure side in a rotationaldownstream position with respect to a coverplate about the engine axis.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the first wall includes a first wallportion with a multiple of apertures.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, an outer rim that extends from the portion.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, a static seal that extends radially inward from theouter rim that extends from the radial portion.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes a knife edge that extends from the coverplate toseal with the static seal.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the outer rim, the radial first wallportion and the first wall define a generally U-shape in cross-section.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the coverplate includes a multiple ofcoverplate apertures to receive the discharge air.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the second wall includes an extendedportion with a multiple of apertures.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the multiple of bypass apertures arecircular.

A gas turbine engine according to another disclosed non-limitingembodiment of the present disclosure includes a coverplate for a turbinerotor defined about an engine longitudinal axis, the coverplateincluding a multiple of coverplate apertures; and an on-board injectorwith a multiple of airfoil shapes between a first wall and a second wallto define an annular inlet about the engine longitudinal axis, themultiple of airfoil shapes operable to segregate and direct dischargeair from the annular inlet toward the multiple of coverplate apertures,the on-board injector including a multiple of bypass apertures eachalong a radial axis transverse to the engine longitudinal axis andthrough each of the multiple of airfoil shapes, the first wall, and thesecond wall.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the on-board injector is a radial on boardinjector.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the on-board injector is an angled on boardinjector.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein each the multiple of airfoil shapes includea pressure side and a suction side, the pressure side in a rotationaldownstream position with respect to a coverplate about the engine axis.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the multiple of airfoil shapes define acascade exit to segregate the discharge air.

A method of managing purge air within a turbo machine according toanother disclosed non-limiting embodiment of the present disclosureincludes a segregating discharge air from an annular inlet with amultiple of airfoil shapes, the annular inlet defined around an enginelongitudinal axis; and directing purge air through a multiple of bypassapertures each along a radial axis transverse to the engine longitudinalaxis and through each of the multiple of airfoil shapes.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, tangentially directing the discharge air.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be appreciated, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiments. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a fragmentary axial cross section of a portion of the turbinesection of a gas turbine engine showing a tangential on-board injector(TOBI) nozzle for the distribution of cooling air;

FIG. 2 is an enlarged axial cross section view of a tangential on-boardinjector (TOBI) used to distribute discharge air for cooling the turbinetaken along line 2-2 in FIG. 3;

FIG. 3 is a partially broken perspective view of the TOBI from anannular inlet perspective;

FIG. 4 is an enlarged axial cross section view of a tangential on-boardinjector (TOBI) used to distribute discharge air for cooling the turbinetaken along line 4-4 in FIG. 5;

FIG. 5 is a partially broken perspective view of the TOBI from a cascadeexit perspective;

FIG. 6 is an enlarged axial cross section view of an angled on-boardinjector (AOBI) used to distribute discharge air for cooling theturbine;

FIG. 7 is a sectional view of the AOBI taken along line 7-7 in FIG. 6;

FIG. 8 is an enlarged axial cross section view of a radial on-boardinjector (ROBI) used to distribute discharge air for cooling theturbine; and

FIG. 9 is a sectional view of the ROBI taken along line 9-9 in FIG. 8.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a portion of a gas turbine engine 10.Although depicted as a turbofan in the disclosed non-limitingembodiment, it should be appreciated that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of turbo machines.

The gas turbine engine 10 generally includes a compressor section 12 anda turbine section 19 mounted on a rotor shaft 15 to form a spool thatrotates about an engine longitudinal axis A. In this disclosednon-limiting embodiment, the turbine 19 is a high pressure turbine. Thecompressor 12 includes a hub 14 mounted to the rotor shaft 15. Adischarge outlet 16 expels discharge air D from the compressor 12 to aturbine inlet 20 via passages 18. A turbine rotor hub 22 that supportsrotor blades 24 is mounted on the shaft 15. The blades 24 receive andexpand the discharge air D from the turbine inlet 20.

Purge air P flow is produced within the compressor section 12, anddirected to the turbine section 19 through a series of passages. Forexample, compressor seals 26 and 28 arranged between the hub 14 andengine housing may leak purge air P into cavities 30 and 31. The purgeair P then leaks past seal 32 and reaches the turbine 19.

An on-board injector 44 which, in this disclosed non-limitingembodiment, is a tangential on-board injector (TOBI) delivers dischargeair D to a space 40 near the turbine 16 for cooling the turbine rotorhub 22. A baffle 43 may be arranged between the passage 18 and theon-board injector 44 to turn the air abruptly to separate debris beforecommunication to the turbine 19. The on-board injector 44 is generallyparallel to the engine longitudinal axis A.

A coverplate 36 separates the on-board injector 44 and the turbine rotorhub 22. A multiple of coverplate apertures 38 are provided in thecoverplate 36 to direct cooling air C from the on-board injector 44 tobe directed into the turbine rotor hub 22.

With reference to FIG. 2, the on-board injector 44 generally includes afirst wall 60, a second wall 62 spaced from the first wall to define anannular inlet 64 about the engine longitudinal axis A, and a multiple ofairfoil shapes 66 between the first wall 60 and the second wall 62 tosegregate discharge air from the annular inlet 64 (also shown in FIG.3). The first and second wall 60, 62 are annular walls defined about theengine axis A. It should be appreciated that the on-board injector 44may be manufactured of separate assembled components or integrallymanufactured such as via an additive manufacturing process.

Each of the multiple of airfoil shapes 66 include a respective bypassaperture 68 each along a radial axis B (FIG. 4) transverse to the enginelongitudinal axis A and the respective first and second wall 60, 62.Each of the multiple of airfoil shapes 66 includes a first sidewall 70that may be convex and defines a suction side, and a second sidewall 72that may be concave and define a pressure side. Sidewalls 70, 72 arejoined at a leading edge 74 and at an axially spaced trailing edge 76.More specifically, each airfoil trailing edge 76 is spaced chordwise anddownstream from the airfoil leading edge 74 to segregate the dischargeair from the annular inlet 64 though a cascade exit 80 (FIG. 5). Thatis, the cascade exit 80 is defined by the sidewalls 70, 72 whichseparate the initially annular flow into the annular inlet 64 such thatthe pressure side is in a rotational downstream position with respect tothe coverplate 36 about the engine longitudinal axis A.

The sidewalls 70, 72 extend radially between the first and second wall60, 62 to segregate the discharge air from the annular inlet 64 and turnthe discharge air in a tangential direction coordinated with arotational direction of the coverplate 36 and the turbine rotor hub 22.In one disclosed non-limiting embodiment, each trailing edge 76 isarranged about 80 degrees to axial. In another disclosed non-limitingembodiment, each trailing edge 76 is arranged about 10 degrees tocircumferential.

The first wall 60 further includes a radial first wall portion 82 with amultiple of apertures 83 in communication with a cooling air supplycavity 84. The radial first wall portion 82 extends into an outer rimportion 86 operable to support a static seal 88. The static seal 88extends radially inward from the outer rim portion 86 to interface witha knife edge 89 that extends from the coverplate 36. That is, the outerrim portion 86, the radial first wall portion 82 and the first wallportion 60 defines a generally U-shape in cross-section.

The second wall 62 includes an extended portion 90 with a multiple ofapertures 92 in communication with the cooling air supply cavity 84. Theapertures 83, 92 are optional and may facilitate, for example, mass flowdistribution between the cooling air supply cavity 84, an outer rimsealing cavity 94, and an inner turbine rotor purge cavity 96. The massflow through aperture 83 is preferably zero. The mass flow throughaperture 92 is minimized with the combined flow from aperture 92 and thepurge mass flow P substantially equal to the mass flow required forpurging an outermost rim cavity 100.

With reference to FIG. 4, the bypass apertures 68 communicate, orbypass, airflow from the inner turbine rotor purge cavity 96 to theouter rim sealing cavity 94 such that the airflow does not cross thedischarge air from the annular inlet 64 that is directed into thecoverplate apertures 38. The bypass apertures 68 may be circular orotherwise shaped such as teardrop or oval to further accommodate and/ormodify airflow therethrough. In one example, 20-40 bypass apertures 68each of about 0.25 inches (6.25 mm) in diameter are provided.

This architecture minimizes or avoids the ejector effect of aconventional cascade exit. The cascade forms a nozzle that swirls andaccelerates the cooling flow to match the rotational velocity of therotor. The increase in momentum of this mass flow can entrainsurrounding air, and pull it into the high velocity flow. Previously,the low momentum purge air P had to cross the plane of the cascade exit.The crossing purge flow P both inhibited the flow of the discharge airfrom the cascade exit and added to the mixing between the cooling flow Cand purge flow P, which raised the temperature of the cooling airreaching the rotor, lowering the cooling air overall momentum, andthereby reducing cooling effectiveness.

The bypass apertures 68 essentially operate as vents through the cascadesuch that the purge mass flow can pass through the “solid walls” createdby the cascade flowpath on-board injector 44, and satisfy the K/E massflow requirements. Thus, the crossing flow is greatly reduced, theon-board injector cooling flow is provided to the rotor with lesspollution, and a lower overall temperature results. In one example, thetemperature is operational reduced by 4-5%. Lower blade cooling airtemperature allows the rotor cooling flow to be reduced for a cycleimprovement, a reduction in TSFC, and improved turbine efficiency.

It should be appreciated that in some cases there will be a contributionfrom the on-board injector 44 discharge flow to form the purge air P. Ifthe turbine rotor cavity is effectively sealed off from the HPCdischarge air, then the on-board injector 44 inlet mass flow at cavity84 is about equal to the cooling flow C, the purge air P, the mass flowthrough the multiple of apertures 83 and the mass flow through aperture92. Further, it may be desired that the mass flow through the multipleof apertures 83 is zero, while the purge air P and the mass flow throughaperture 92 pass through the bypass apertures 68.

With reference to FIG. 6, in another disclosed non-limiting embodiment,the on-board injector 44A is an angled on-board injector (AOBI). Theon-board injector 44A is as described above but angled with respect tothe engine longitudinal axis A (also shown in FIG. 7).

With reference to FIG. 8, in another disclosed non-limiting embodiment,the on-board injector 44B is a radial on-board injector (ROBI). Theon-board injector 44B is as described above but generally perpendicularto the engine longitudinal axis A (also shown in FIG. 9). It should beappreciated that other arrangements will benefit herefrom.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be appreciated that steps may be performed in any order,separated or combined unless otherwise indicated and will still benefitfrom the present disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beappreciated that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed:
 1. A system for a gas turbine engine comprising: acoverplate for a turbine rotor defined about an engine longitudinalaxis, said coverplate including a multiple of coverplate apertures; andan on-board injector with a multiple of airfoil shapes between a firstwall and a second wall to define an annular inlet about the enginelongitudinal axis, said multiple of airfoil shapes operable to segregateand direct discharge air from the annular inlet toward said multiple ofcoverplate apertures, said on-board injector including a multiple ofbypass apertures each along a radial axis transverse to the enginelongitudinal axis, one of each of said multiple of apertures extendsthrough one of said multiple of airfoil shapes, said first wall, andsaid second wall, wherein said on-board injector is an axial on-boardinjector.
 2. The system as recited in claim 1, wherein each of saidmultiple of airfoil shapes include a pressure side and a suction side,said pressure side in a rotational downstream position with respect tosaid coverplate about said engine axis.